High pressure ratio gas turbine engine

ABSTRACT

A gas turbine engine (10) comprising:a high pressure turbine (17);a low pressure turbine (19);a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27);a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four or five compressor stages (14);the high pressure compressor (15) consists of eight or nine compressor stages;the low pressure turbine (19) comprises four or more stages; andthe high pressure compressor (15) and low pressure compressor (14) together define a core overall pressure ratio of greater than 36:1.

The present disclosure relates to gas turbine engine for an aircraft

Existing gas turbine engines are known, in which a reduction gearbox isprovided between a turbine and a propulsive fan. Such engines are knownas “geared turbofans”.

It desirable to increase the fuel efficiency of such engines. There areessentially two methods to increase fuel efficiency—increased propulsiveefficiency, and increased thermal efficiency. It is an objective of thepresent invention to provide a gas turbine engine architecture thatprovides improved fuel efficiency

According to a first aspect there is provided a gas turbine enginecomprising:

-   a high pressure turbine;-   a low pressure turbine;-   a high pressure compressor coupled to the high pressure turbine by a    high pressure shaft;-   a propulsor and a low pressure compressor coupled to the low    pressure turbine via a low pressure shaft and a reduction gearbox;    wherein-   the low pressure compressor consists of four or five compressor    stages,-   the high pressure compressor consists of eight or nine compressor    stages;-   the low pressure turbine comprises four or more stages; and-   the high pressure compressor and low pressure compressor together    define a core overall pressure ratio at cruise of greater than 36:1.

The inventors have found that the above defined characteristics canprovide a gas turbine engine having a high overall pressure ratio (andso high thermal efficiency), with relatively few stages, and with arelatively low pressure ratio high pressure compressor. Such arelatively low pressure ratio high pressure compressor can provide fornumerous advantages, such as a reduction in variable stator stages andbleed valves, which can in turn result in reduced weight and cost.

The core overall pressure ratio may be between 36:1 and 60:1. Theoverall pressure ratio may be any of 36:1, 38:1, 40:1, 42:1, 44:1, 46:1,48:1, 50:1, 52:1, 54:1, 56:1 or 58:1.

The low pressure compressor may define a cruise average stage pressureratio of between 1.24:1 and 1.35:1.

The low pressure compressor may define a cruise pressure ratio ofbetween 2.3 and 4.5.

The high pressure compressor may define a cruise pressure ratio ofbetween 8:1 and 18:1.

The high pressure compressor may define a cruise average pressure ratioof between 1.3 and 1.42.

The high pressure turbine may consist of two or fewer stages.

The low pressure turbine may consist of four stages.

The low pressure compressor may be positioned axially upstream of thehigh pressure compressor. The high pressure compressor may be arrangedto receive (for example directly receive, for example via a generallyannular duct) flow from the low pressure compressor.

The gearbox may be arranged to be driven by the shaft that is configuredto rotate (for example in use) at the lowest rotational speed (forexample the low pressure shaft in the example above).

The engine may comprise a core casing and a nacelle, wherein at leastone of the core casing the nacelle comprise carbon composite material.

Any type of reduction gearbox may be used. For example, the gearbox maybe a “planetary” or “star” gearbox, as described in more detailelsewhere herein. The gearbox may have any desired reduction ratio(defined as the rotational speed of the input shaft divided by therotational speed of the output shaft), for example greater than 2.5, forexample in the range of from 3 to 4.2, or 3.2 to 3.8, for example on theorder of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4,4.1 or 4.2. The gear ratio may be, for example, between any two of thevalues in the previous sentence. Purely by way of example, the gearboxmay be a “star” gearbox having a ratio in the range of from 3.1 or 3.2to 3.8. In some arrangements, the gear ratio may be outside theseranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the propulsor andcompressors. For example, the combustor may be directly downstream of(for example at the exit of) the high pressure compressor. By way offurther example, the flow at the exit to the combustor may be providedto the inlet of the high pressure turbine.

The propulsor may be in the form of an open rotor, or a ducted fan.

Each compressor and/or turbine stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform. The radius of the fan may be measuredbetween the engine centreline and the tip of a fan blade at its leadingedge. The fan diameter (which may simply be twice the radius of the fan)may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm,250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm(around 110 inches), 290 cm (around 115 inches), 300 cm (around 120inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches),340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm(around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165inches). The fan diameter may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 240 cm to 280cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allvalues being dimensionless). The fan tip loading may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds), for example in the range offrom 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall core pressure ratio of a gas turbine engine as describedand/or claimed herein may be defined as the ratio of the stagnationpressure downstream of the fan to the stagnation pressure at the exit ofthe highest pressure compressor (before entry into the combustor). Byway of non-limitative example, the overall pressure ratio of a gasturbine engine as described and/or claimed herein at cruise may begreater than (or on the order of) any of the following: 36, 40, 45, 50,55. The overall core pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400 K, 1450 K, 1500 K,1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800 K to 1950 K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN. The fan of a gas turbine as described and/orclaimed herein may have any desired number of fan blades, for example14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of, the gas turbine engine that providesa thrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The method may comprise, at cruise conditions, operating the lowpressure compressor (14) to provide a pressure ratio of between 2.4:1and 3.3:1, operating the high pressure compressor (15) to provide apressure ratio of less than 18:1, and operating the low and highpressure compressors (14, 15) to provide a pressure ratio of less than36:1.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of thegas turbine engine of FIG. 1;

FIG. 3 is a close up sectional side view of a turbine section of the gasturbine engine of FIG. 1;

FIG. 4 is a sectional front view of a reduction gearbox of the gasturbine engine of FIG. 1;

FIG. 5 is a graph illustrating a design space for the compressor sectionof FIG. 2;

FIG. 6 is a graph illustrating a design space for high pressurecompressors of gas turbine engines in accordance with the presentdisclosure; and

FIG. 7 is a graph illustrating a design space for low pressurecompressors of gas turbine engines in accordance with the presentdisclosure.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. In some cases, the low pressure compressor 14may be referred to as an intermediate pressure compressor (IPC).Similarly, the low pressure turbine may be referred to as anintermediate pressure turbine (IPT). A nacelle 21 surrounds the gasturbine engine 10 and defines a bypass duct 22 and a bypass exhaustnozzle 18. The bypass airflow B flows through the bypass duct 22.

The fan 23 is attached to and driven by the low pressure turbine 19 viaa shaft 26 and an epicyclic gearbox 30.

The engine core 11 is surrounded by a core casing 37, which contains thecompressor 14, 15, combustor 16 and turbines 17, 19. The core casing 37comprises one or more handling bleeds comprising one or more valves 38configured to communicate between the core compressor flow path A (e.g.at the downstream end of the high pressure compressor 15) and the fanflowpath B. The core casing 37 comprises a carbon composite materialsuch as carbon fibre reinforce plastic (CFRP).

Similarly, the engine nacelle 21 comprises a carbon composite material,such as CFRP. For example, a Thrust Reverser Unit (TRU) 39 provided at arear of the nacelle 21 may comprise carbon composite material.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 4. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 is statically mounted, whichconstrains the planet gears 32 whilst enabling each planet gear 32 torotate about its own axis. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that that iscoupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Such an arrangement is typically referred to asa “star” gearbox.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 4. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 4. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIG. 4 is ofthe planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIG. 4 is by way ofexample only, and various alternatives are within the scope of thepresent disclosure. Purely by way of example, any suitable arrangementmay be used for locating the gearbox 30 in the engine 10 and/or forconnecting the gearbox 30 to the engine 10. By way of further example,the connections (such as the linkages 36, 40 in the FIG. 1 example)between the gearbox 30 and other parts of the engine 10 (such as theinput shaft 26, the output shaft and the fixed structure 24) may haveany desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. By way of further example, the gasturbine engine shown in FIG. 1 has a split flow nozzle 18, 20 meaningthat the flow through the bypass duct 22 has its own nozzle 18 that isseparate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

Referring now to FIG. 2, the low pressure and high pressure compressor14, 15 are shown in more detail. As can be seen, each of the compressorscomprises a multi-stage, axial flow compressor.

The low pressure compressor consists of four or five stages (i.e. nomore than five stages, and no fewer than four stages) 41 a-d. Each stage41 a-d comprises at least one respective compressor rotor 43, and maycomprise a respective stator 44. The respective rotor 43 and stator 44are generally axially spaced. In the present case, the first stator 44is downstream in core flow of the first rotor 43. One or more furtherstators such as an inlet stator (not shown) may be provided—however,since no additional rotor is associated with the inlet stator, this doesnot constitute an additional stage, since no pressure rise is providedby the inlet stator alone. As will be appreciated by the person skilledin the art, the rotors 43 are coupled to the respective shaft (i.e. thelow pressure shaft 26 in the case of the low pressure compressor 14) bycorresponding discs 46 a-d, and so turn with the shaft 26. On the otherhand, the stators 44 are held stationary. In some cases, the stators 44may pivot about their long axes, to adjust the angle of attack and inletand outlet area for the respective compressor stage. Such stators areknown as “variable stator vanes” or VSVs.

The high pressure compressor 15 similarly consists of either eight ornine stages, and in the described embodiment consists of eight stages.Again, each stage comprises at least a rotor, and may also comprise astator.

The turbine is shown in FIG. 3. To drive the high pressure compressor15, a high pressure turbine 17 having two stages 47 a, 47 b may benecessary. Again, the number of turbine stages can be determined in asimilar manner to the number of compressor stages. Alternatively, asingle turbine stage may be provided. In particular, it has been foundthat high pressure compressors having cruise pressure ratios up to 13:1can be driven by single stage turbines.

Similarly, to drive the low pressure compressor 14 and fan 23, at leastfour low pressure turbine stages 49 a-d are provided. In some cases,five low pressure turbine stages may be provided.

Between them, the high and low pressure compressors 15, 16 define amaximum in use overall core pressure ratio (OPR). The core OPR isdefined as the ratio of the stagnation pressure upstream of the firststage 44 of the low pressure compressor 15 to the stagnation pressure atthe exit of the highest pressure compressor 16 (before entry into thecombustor). The core OPR excludes any pressure rise generated by the fan23 where the fan provides air flow to the core, so a total engineoverall pressure ratio (EPR) may be higher than the core OPR. In thepresent disclosure, the overall core OPR is at least 36:1, and may bebetween 36:1 and 56:1. In the described embodiment, the core OPR is 40,and may take any value between these upper and lower bounds. Forexample, the core OPR may be any of 36, 40, 45, 50, 55 and 60, or evenhigher.

As will be understood, the core OPR will vary according to atmospheric,flight and engine conditions. However, the cruise OPR is as definedabove.

As will be understood, a large design space must be considered whendesigning a gas turbine engine to determine an optimal engine withrespect to a chosen metric (such as engine weight, cost, thermalefficiency, propulsive efficiency, or a balance of these). In manycases, there may be a large number of feasible solutions for a given setof conditions to achieve a desired metric.

One such variable is core OPR. As core OPR increases, thermal efficiencyalso tends to increase, and so a high OPR is desirable. Even once aparticular OPR is chosen however, a number of design variables must bechosen to meet the chosen OPR.

One such design variable is the amount of pressure rise provided by thelow pressure compressor 15 relative to that provided by the highpressure compressor 16 (sometimes referred to as “worksplit”). As willbe understood, the total core OPR can be determined by multiplying thelow pressure compressor pressure ratio (i.e. the ratio between thestagnation pressure at the outlet of the low pressure compressor to thestagnation pressure at the inlet of the low pressure compressor 15) bythe high pressure compressor ratio (i.e. the ratio between thestagnation pressure at the outlet of the high pressure compressor 16 tothe stagnation pressure at the inlet of the high pressure compressor16). Consequently, a higher core OPR can be provided by increasing thehigh pressure compressor ratio, the low pressure compressor ratio, orboth.

The inventors have found that a particularly efficient work split for agas turbine engine having a core OPR in the above described range can beprovided by providing a low pressure compressor 14 consisting of four orfive stages, and having a pressure ratio of between 2.4:1 and 4.5:1, anda high pressure compressor is then provided having a pressure ratioabove 8:1, such that the overall core pressure ratio is above 36:1. Ithas been found to be feasible to provide a pressure ratio of 18:1 on ahigh pressure compressor provided on a single shaft using currenttechnology using a reasonable number of compressor stages, withoutrequiring an excessive number of variable stages, and at a reasonablerotational speed to give high overall efficiency. Consequently, toprovide the necessary core OPR, a low pressure compressor ratio ofbetween 2.4:1 and 4.5:1 is required.

Similarly, there are several ways to increase the compressor pressureratio. A first method is to increase the stage loading. Stage loading isdefined as the stagnation pressure ratio across an individual stage(rotor and stator) of a compressor. Similarly, an average stage loadingcan be defined as the sum of the stage loadings of each compressor stageof a compressor, divided by the number of stages. For example, in thepresent disclosure, the average stage loading of the low pressurecompressor 14 is between 1.24 and 1.35. This can in turn be managed byone or more of increasing the rotor speed at the maximum compressionconditions, increasing the turning provided by the blades, or increasingthe radius of the tips of the compressor rotors, which in turnnecessitates an increase in the radius of the roots of the compressorrotors to maintain a given flow area. Each of these options hasassociated advantages and disadvantages. For instance, increasing lowpressure compressor rotor speed necessitates either an increase in thereduction ratio of the gearbox 30, or a reduction in the fan 23 radius,in order to maintain fan tip speeds at a desired level for noise andefficiency reasons. On the other hand, increasing the compressor tipradius necessitates an increase in weight, in view of the largercompressor discs that are required. Increased turning of the airflow mayresult in lower surge margin, and reduced efficiency. In any case, ahigher stage loading may result in a lower efficiency, since theincreased rotor tip speed or higher turning leads to lower compressorefficiencies, in view of losses associated with aerodynamic shocks asthe tips significantly exceed the speed of sound.

A second option is to increase the number of stages in the respectivecompressors, thereby maintaining a low stage loading, low rotationalspeed, and low disc weight. Again, this can be achieved by adding astage to either the low pressure compressor 15 or high pressurecompressor 16. However, this will generally result in a higher weightand cost associated with the additional stage.

A further complication is the presence of the gearbox 30. The gearboxprovides additional design freedom, since, as noted above, the gearboxreduction ratio can be selected to provide a preferred fan tip speedindependently of both fan radius and low pressure compressor rotorspeed. However, the gearbox also presents constraints in view of itslarge size. Consequently, the large radius required radially inward ofthe fan 23 inherent in a geared turbofan having an epicyclic gearboxdictates a fan 23 having a large hub radius, i.e. a large radialdistance between the engine centre 9 and the aerodynamic root of the fanblades 23. Furthermore, in view of the relatively slow turning fantypical of geared turbofans, relatively little pressure rise is providedby the inner radius of the fan 23, and so geared turbofans tend to havea high hub to tip ratio fan 23.

The inventors have explored this design space, and found an optimumrange of stage numbers and compressor pressure ratios, that provides anoptimal mix of weight and efficiency.

In the present disclosure, a relatively large pressure ratio is providedby the low pressure compressor 14, compared to prior geared turbofans.This is achieved using a four or five stage compressor 14, which rotatesat relatively low speed. This low speed allows a relatively low ratiogearbox to be employed (typically of the order of 3:1), whilemaintaining a relatively low fan blade tip speed, with a high diameterfan. The permits a “star” gearbox, as described above, in conjunctionwith a high bypass ratio (typically greater than 10). Such anarrangement has several advantages over planetary gearboxes (in whichthe ring gear is held static, and the planet carrier is used to drivethe fan), such as more convenient oil distribution.

In view of the relatively high pressure high work provided by the lowpressure compressor 14, additional work is required from the lowpressure turbine 19. In view of the relatively low shaft speed, theinventors have found that a fourth or even a fifth turbine stage isnecessary, which is unusual on a geared turbofan. This is because thelow rotational speed results in low turbine tip speeds, which results inrelatively low work per revolution. Consequently, a four or five stagelow pressure turbine is provided in this case.

The above optimum parameters define a design space for the compressor,as illustrated in FIGS. 6 and 7.

One corner of the design space is defined by the maximum low pressurecompressor 14 cruise pressure ratio (4.5:1), and the minimum highpressure compressor 15 cruise pressure ratio (8:1) to achieve theminimum require overall core pressure ratio (36:1). Above this lowpressure cruise pressure ratio (4.5:1), it has been found thatcompressor stability cannot be assured, without increasing eitherrotational speed or diameter (and so compressor blade tip speed ineither case), or increasing low pressure compressor stage count abovefive stages. However, where compressor tip speed is increased,efficiency begins to fall, and so the advantages of higher loading arelost. The inventors have found that a low pressure cruise pressure ratioof 4.5:1 can be provided with no more than five stages. Indeed, theinventors have found that a cruise low pressure ratio of up to 3.3:1 canbe provided with only four low pressure compressor stages. Similarly,overall engine efficiency suffers when the overall core engine pressureratio falls below 36:1, particularly in view of the relatively smallpressure rise generated by the fan in a geared turbofan.

A second corner of the design space is defined by the maximum lowpressure compressor 14 cruise pressure ratio (4.5:1), and the maximumhigh pressure compressor 15 cruise pressure ratio (18:1) that canreasonably be sustained, without requiring excessive stage numbers, andincreased weight. This combination gives an overall core pressure ratioof 80:1. Above this value, increases in thermal efficiency begin to beoutweighed by increases in weight, and so the design goals of increasedoverall propulsion system efficiency are not achieved. In particular,the inventors have found that the above parameters can be provided usinga high pressure compressor having eight or nine stages, with relativelylow work per stage. This relatively low work per stage provides for highcompressor efficiency, while the high overall pressure ratio results inhigh overall engine efficiency.

A third corner of the design space is defined by the maximum highpressure compressor 15 cruise pressure ratio (18:1) that can reasonablybe sustained, and the minimum low pressure compressor cruise pressureratio (2.4:1) that requires four compressor stages. Below this value,only three low pressure compressor stages are required, and so theengine weight is not optimised. This combination gives an overall corepressure ratio of 43:1, which provides good thermal efficiency, with asmall number of overall compressor stages.

A fourth corner of the design space is defined by the minimum highpressure compressor cruise pressure ratio required to achieve therequired overall core pressure ratio of 36:1 at the minimum low pressurecompressor 14 stage loading for which four compressor stages arerequired (2.4:1). This gives a high pressure compressor cruise pressureratio of approximately 15:1.

A fifth corner of the design space is defined. At this point, a higherlow pressure compressor cruise pressure ratio of 3.0:1 is provided, anda lower high pressure compressor cruise pressure ratio of 12:1 isprovided, while providing the minimum overall core compressor pressureratio of 36:1.

In one example, such as that shown in FIG. 2, the engine comprises a lowpressure compressor consisting of four stages, and a high pressurecompressor consisting of eight stages. FIG. 5 illustrates the designspace for this combination. As can be seen, the high pressure compressoroverall pressure ratio can vary between 15 and 17:1, and the lowpressure compressor pressure ratio can vary between 2.7:1 and 3.3:1.Such a combination is thought to allow for a minimum total stage count,while providing a high overall pressure ratio, since the loading of bothspools is maximised without compromising stability.

The designer is hence taught how to design a compressor which achievesthe desired characteristics of high overall core cruise pressure ratio(greater than 36:1), while minimising stage count and maximisingcompressor efficiency.

Two example gas turbine engines that have been considered by theinventors are described below.

A first example engine has a maximum take-off thrust at sea level underISO conditions of approximately 45,000 pounds-force (lbf). The lowpressure compressor has four stages, and is configured to provide acruise pressure ratio of approximately 2.8:1. The high pressurecompressor is configured to provide a cruise pressure ratio ofapproximately 13:1. This gives an overall core pressure ratio ofapproximately 36:1. Such an engine is thought to provide an optimum mixof weight and thermal efficiency for an engine in this class, sinceweight is a more important factor in this class than for higher thrustengines, in view of the shorter typical mission ranges of aircraft forwhich engines of this thrust are designed.

A second example engine has a maximum take-off thrust at sea level underISO conditions of approximately 84,000 pounds-force (lbf). The lowpressure compressor has four stages, and is configured to provide acruise pressure ratio of approximately 2.8:1. The high pressurecompressor is configured to provide a cruise pressure ratio ofapproximately 17:1. This gives an overall core pressure ratio ofapproximately 48:1. Such an engine is thought to provide an optimum mixof weight and thermal efficiency for an engine in this class, sincethermal efficiency is a more important factor in this class than forlower thrust engines, in view of the longer typical mission ranges ofaircraft for which engines of this thrust are designed.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine comprising: a high pressure turbine; a lowpressure turbine; a high pressure compressor coupled to the highpressure turbine by a high pressure shaft; a propulsor and a lowpressure compressor coupled to the low pressure turbine via a lowpressure shaft and a reduction gearbox; wherein the low pressurecompressor consists of four or five compressor stages; the high pressurecompressor consists of eight or nine compressor stages; the low pressureturbine comprises four or more stages; and the high pressure compressorand low pressure compressor together define a core overall pressureratio of greater than 36:1.
 2. A gas turbine engine according to claim1, wherein the core overall pressure ratio is between 36:1 and 60:1. 3.A gas turbine engine according to claim 1, wherein the low pressurecompressor defines a cruise average stage pressure ratio of between1.24:1 and 1.35:1.
 4. A gas turbine engine according to claim 1, whereinthe low pressure compressor defines a cruise pressure ratio of between2.3 and 4.5.
 5. A gas turbine engine according to claim 1, wherein thehigh pressure compressor defines a cruise pressure ratio of between 8:1and 18:1.
 6. A gas turbine engine according to claim 1, wherein the highpressure compressor defines a cruise average pressure ratio of between1.3 and 1.42.
 7. A gas turbine engine according to claim 1, wherein thehigh pressure turbine consists of two or fewer stages.
 8. A gas turbineengine according to claim 1, wherein the low pressure turbine consistsof four stages.
 9. A gas turbine engine according to claim 1, whereinthe reduction gearbox comprises a star gearbox.
 10. A gas turbine engineaccording to claim 1, wherein the reduction gearbox defines a reductionratio of approximately
 3. 11. A gas turbine engine according to claim 1,wherein the low pressure compressor is positioned axially upstream ofthe high pressure compressor.
 12. A gas turbine engine according toclaim 1, wherein the propulsor is in the form of an open rotor, or aducted fan.
 13. A gas turbine engine according to claim 1, wherein eachcompressor and/or turbine stage comprises a row of rotor blades and arow of stator vanes, which may be variable stator vanes.
 14. A method ofoperating the gas turbine engine of claim 1, comprising, at cruiseconditions, operating the low and high pressure compressors to provide apressure ratio of greater than 36:1.